Integrated and match machined rocket nozzle and process of making same



Dec. 31, 1968 H. c. EMERSON v INTEGRATED AND MATCH MACHINED ROCKETNOZZLE Sheet AND PROCESS OF MAKING SAME Filed April 19, 1965 INVENTOR.

H. c. EMERSON ATTORNEY Dec. 31, 1968 H. c. EMERSON 3,418,707

INTEGRATED AND MATCH MACHINE!) ROCKET NOZZLE AND PROCESS OF MAKING SAMEFiled April 19, 1965 Sheet g of 6 25 INVENTOR. Fl 7 H. C. EMERSONATTORNEY Dec. 31, 1968 H. EME ON 3,418,707

INTEGRATED AND M H MAC ED ROCKET NOZZLE AND PROCESS OF MAKING SAME FiledApril 19. 1965 Sheet ,3 of e INVENTOR.

H. C. EMERSON ATTORNEY Dec. 31, 1968 H. c. EMERSON 3,418,707

INTEGRATED ND MATCH MACHINE!) ROCKET NOZZLE Filed April l9. 1965INVENTOR.

H. C. EMERSON FIG. 12

fun.

ATTORNEY 1968 H. c. EMERSON 3,418,707

INTEGRATED AND MATCH MACHINE!) ROCKET NOZZLE AND PROCESS 0E )M KING SAMEFiled April 19 1965 H ATTORNEY Sheet 5 'of6 INVENTOR.

H. C. EMERSON I Dec. 31, 1968 H. c. EMERSON INTEGRATED AND MATCHMACHINED ROCKET NOZZLE AND PROCESS OF MAKING SAME Sheet Filed April 19,1965 INVENTOR.

H. C. EMERSON ATTORNEY United States Patent INTEGRATED AND MATCHMACHINED ROCKET NOZZLE AND PROCESS OF MAKING SAME Herfi C. Emerson,Chula Vista, Calif., assignor to Rohr Corporation, San Diego, Calif., acorporation of California Filed Apr. 19, 1965, Ser. No. 449,076 4Claims. (Cl. 29-157) ABSTRACT OF THE DISCLOSURE A rocket nozzleshell-ablative liner composite is disclosed in which a large hot sized,high strength unitary nozzle is formed of annularly weldedfrusto-conical ring sections of varying diameters and cone angles. Eachof the ring sections is formed of arcuate ring segments welded togetherto form a unitary frusto-conical ring section, and each ring segment iscut and contoured plate metal. The Welded ring sections and unitarynozzle are hot sized to remove distortion, and the external surfaces ofthe resulting nozzle structure is only nominally machined to designconfiguration. The internal surface dimensions of the nozzle aremeasured numerically to receive a match machined ablative liner which isbonded thereto. The liner has an inner layer of an ablative material andan outer layer of a resin impregnated fibre glass fabric which ismachined to the inner dimensions of the nozzle shell.

Background This invention relates to rocket nozzles in general and morespecifically to a segmented rocket nozzle shell-ablative liner compositeand the process of integrating the shell and match machining the linerthereto.

The rocket nozzle shell-ablative liner composite and the processconsidered herein allows for the fabrication of large rocket nozzleshell-ablative liner composites previously considered impractical, ifnot impossible, due to limitations imposed by size and economic factors.Certain of the prior art nozzle-ablative composites that heretofore havebeen fabricated incorporate a single forging for the shell. This priorart method is necessarily limited to nozzle-ablative composites ofrelatively small size. The use of a single forging, moreover,necessarily requires additional material to form the forging envelopethat subsequently has to be machined 011.

Other prior art nozzle shell-ablative liner composites of larger sizesalso have heretofore required a forging or a combination of cone sectionforgings joined together by of the smaller nozzle shell-ablativecomposites, excessive material is required for the forging envelopewhich is subsequently machined away.

Present invention In accordance with the present invention, whereinplate segments are used, only a nominal excess of material for clean-upmachining purposes of the integrated nozzle shell is required, therebygreatly reducing the costly prior art machining process necessarilyincurred in the use of forgings. By using plate segments with only anominal excess of material for the larger size nozzle shell-ablativecomposites, moreover, even greater savings are realized. The size of thecone sections of the larger nozzle shell-ablative composites accordingto the present method are thus not limited to the size of the forgingsas in the prior art methods but only to the size of plate stockavailable.

By match machining the ablative material to mate with the internalsurfaces of the nozzle shell as contemplated in the present invention,the handling and fit-up prob lems attendant with large size nozzleshells are greatly minimized. Using numerically controlled machiningtechniques to form the liner to the size and shape of the internalsurface of the nozzle shell, the ablative material can be fitted andbonded to the nozzle shell without requiring an excessively thickadhesive line.

Objects An object of the present invention is to provide a fabricationmethod for rocket nozzle shell-ablative liner composites of a variety ofsizes and materials.

Another object is to fabricate nozzle shell-ablative compositespreviously limited in size by using integrated cone ring segments andsections that are joined by a welding process.

Another object is to eliminate the costly machining of excess forgingmaterial by using plate segments containing only a nominal amount ofexcess material.

Another object is to greatly reduce distortion introduced by thefabrication and welding processes by heat treating and sizing theindividual segmented cone sections and the entire nozzle shell onspecialized sizing fixtures.

Another object is to greatly reduce any distortion introduced by thefabrication and Welding processes such that the internal surfaces of thenozzle shell will require no machining or only a minimum of machining inorder to have an ablative liner attached thereto.

Still another object is to match machine the external surfaces of theablative liner material to mate with the internal surfaces of nozzleshell thus eliminating the fitup and handling problems associated withprior methods.

Yet another object is to use numerically controlled machining techniquesto form the external surfaces of the ablative liner to the size andshape of the internal surfaces of the nozzle shell to thereby allow theablative liner material to be fitted and bonded to the nozzle shellwithout requiring an excessively thick adhesive line.

A still further object is to use numerically controlled machiningtechniques to obtain maximum dimensional control of all detail parts andfinal assembly operations for all machining, grinding, tape wrapping andfilament winding.

Still other objects, features and advantages of the present inventionwill become more clearly apparent as the description proceeds, referencebeing had, to the accompanying drawings wherein:

Brief description of the drawings FIG. 1 is a perspective view of acompleted rocket nozzle constructed in accordance with the presentinvention and attached to a rocket motor case;

FIG. 2 is an elevational view of a completed rocket nozzle detached fromthe rocket motor case;

FIG. 3 is a perspective view showing the layout of the plate segments;

FIG. 4 is a fragmentary view illustrating the cutting of a platesegment;

FIG. 5 is a view in perspective of the hot form die arrangement forcontour forming the plate segments;

FIG. 6 is a perspective view showing the machining operation of theplate segment weld preparation;

FIG. 7 is a perspective view illustrating the way in which the platesegments are fit together on a jig prior to welding;

FIG. 8 is a perspective view of the OD. weld jig sections employed, eachsection being used to position each individual cone section;

FIG. 9 is a perspective view illustrating the OD. cone section weldingoperation using one of the OD. weld jig sections shown in FIG. 8;

FIG. 10 is a perspective view illustrating the ID. cone section weldingoperation;

FIG. 11 is a perspective view, partially cut-away, of a hot sizingfixture used for sizing one of the individual cone sections;

FIG. 12 is a persepctive view illustrating the machining operation ofthe weld preparation for the cone section circumferential welds;

FIG. 13 is a perspective view illustrating the manner in which the conesections are positioned prior to welding;

FIG. 14 is a perspective view showing the circumferential weldingoperation;

FIG. 15 is a view of the hot sizing fixture used for sizing the entirenozzle shell;

FIG. 16 is a sectional view, somewhat enlarged, of the hot sizingfixture taken along the line 16-16 of FIG. 15;

FIG. 17 shows an arrangement for machining the external surfaces of thesized nozzle shell;

FIG. 18 is a perspective view showing an arrangement for reading andtape recording the shell dimensions to be used in the numerical controlmachining of the mating ablative liner components;

FIG. 19 is a perspective view, partially cut-away, illustrating thenumerical control machining of the mating ablative liner components;

FIG. 20 is a composite sectional view showing the mating of the ablativeliner components to the internal surface of the nozzle shell; and

FIG. 21 is a composite sectional view showing the wall construction of acompleted rocket nozzle and showing the various materials used in itsconstruction.

Specification Reference is now directed to the drawings for a morecomplete understanding of the invention and first more particularly toFIG. 1 which depicts a fragmentary portion of a rocket of a typedesigned for space exploration. Such a rocket employs a motor case 8 towhich is attached a nozzle 9. Such rockets, including the nose cone andintermediate sections may have overall dimensions of the order of 260inches in diameter and 100 feet in length, and are capable of producinga thrust upwards of three million pounds.

The invention per se is directed to the nozzle 9 which is shown indetail in FIG. 20 to which attention is now directed. It will be seenthat rocket nozzle 9 comprises a shell 10 of a varying converging anddiverging configuration and ablative liner sections 50, 51, and 52assembled and bonded thereto. The shell 10 being of a high strengthmaterial such as 18% nickel maraging steel, serves as the mainstructural element of the nozzle. Accordingly, the shell 10 acts torestrain the thrust force that the nozzle is subjected to as a result ofthe burning of the rocket fuel. The ablative liner material in sections50, 51 and 52 serves to protect the shell 10 in a two-fold manner.First, the major portion of the exhaust gas heating from the burning ofthe rocket fuel is absorbed by the material which is ablated away, and,secondly, the material which does not ablate away insulates the shell10' by absorbing any conducted heat.

Fiberglass material 12 is wrapped around the ablative material ofsections 50, 51 and 52 to provide strength and additional insulation forthe shell 10. A bonding material (not shown) is applied as at 13 betweenthe surfaces of the ablative material sections and of the shell 10 andserves as the means of attaching the ablative sections to the shell.

Attention is now directed to FIGS. 3 and 4 for a detailed description ofthe process by which the rocket nozzle shell is fabricated. To obtainthe plate segments 20, FIG. 4, that make up the cone sections of theshell, flat pattern templets 21 are laid out and scribed on flat platestock 22 of the material used for the shell. Once the segments arescribed on the flat plate stock the flat pattern templets 21 are removedand the segments 20 are cut out of the plate, as shown in FIG. 4, by asuitable cutting process such as plasma are cutting.

Once the flat plate segments 20 are cut from the fiat plate stock 22they are heated to an appropriate annealing temperature such as 1650 F.for 18% nickel maraging steel. While at the appropriate annealingtemperature, each segment is hot formed to the required contour in thehydropress 23, FIG. 5. This contour forming operation can also be formedby cold working the flat plate segments 20 on a press.

When the contour forming operation is completed, each formed platesegment 20 is set up in a milling machine 24, FIG. 6, and a weldpreparation is machined into it. In the instant case where the platesegments are of a substantial thickness, a weld preparation in the formof matched edge grooves referred to in the art as a U-joint, is employedto facilitate the welding process, it being understood where relativelythin plate segments are used a weld preparation may not be required. Instill other instances a similar weld preparation on the oppositesurfaces of the plate segments to be joined may be required.

In FIG. 7 the necessary number of plate segments 20 required to make upa cone section 25 are shown fit together on the weld jig 26, for tackingwelding of the same together prior to making the longitudinal welds 27of the cone section 25. The welding occurs as depicted in FIG. 9 whichshows the making of an OD. weld pass of a longitudinal weld 27 by anautomatic welding machine 29 such as a Linde Unionmelt Welding Head.After all of the OD. weld passes of the longitudinal Welds 27 arecompleted, the cone section 25 is repositioned as shown in FIG. 10 andan ID. weld pass of a longitudinal weld 27 is made thus completing theweld. The nature of large rocket nozzles is such that they requirematerial gages which are too thick to permit single pass welding. Inorder to weld the various thick plate segments a 2 pass OD. and ID.welding technique has been found very successful.

FIG. 11 illustrates a typical hot sizing fixture 31 which is used togreatly reduce any distortion introduced by the welding process in anyone of the cone sections. The cone section is heated to theannealing'temperature of the material and then placed on the sizingfixture while at that temperature. A cover plate 32 is placed over thecone section and force is applied to the cone section by means of thewedges 33, each of which is driven through an opening in a positionerpost 34 therefor. By forcing the cone section 25 down against the rigidcone member 35 of sizing fixture 31, the cone section will be rounded upto the diameter of the cone member and will be properly sized. Once thesizing operation is completed, the cone section 25 will be positioned asshown in FIG. 12 and will have a circumferential weld preparationmachined into it by a boring mill 36.

FIG. 13 shows the manner in which the cone sec tion 25 is fit togetherwith another cone section 37, on the weld jig 28 for tack welding of thesame together prior to making the OD. pass circumferential welds 38 byan automatic welding machine 29 shown in FIG. 14, thereby joining conesection 25 to cone section 37. After all of the OD. pass circumferentialwelds 38 are completed, the shell is repositioned much in the samemanner as in FIG. 10, and the LB. pass circumferential welds are made bythe same automatic welding machine 29.

FIG. 16 illustrates the hot sizing fixture 39 used to reduce thedistortion that has resulted from the circumferential welding process.After the welding operations are complete the shell 10 is heated to theappropriate annealing temperature and placed on the sizing fixture 39 asquickly as possible. The cover plate 40 is then placed over the shell 10and force is applied to the shell 10 'by means of wedges 4, each ofwhich is driven through an opening in the positioner post 42 individualthereto.

The sizing plates 43 are adjusted in such a manner that the combinationof the shrinkage of the shell and the imposed restraint of the sizingfixture 39 will allow the shell to be properly rounded and sized forsubsequent operations.

If a final heat treating operation, such as a 900 F. aging cycle for 18%nickel maraging steel, is required, it is performed after the hot sizingoperation.

After the sizing or heat treating operation is completed the externalsurfaces of the shell 10 are machined on the boring mill 36, as shown inFIG. 17, to satisfy design requirements.

FIG. 18 shows the manner in which the internal surface dimensions of theshell 10 are tape recorded by the recording device 44 to provide arecord for use in the numerically controlled machining of the matingablative liner sections 50, 51 and 52. These dimensions can also beobtained by conventional manual inspection techniques and thentransposed onto tape for the numerically controlled machining of theablative liner sections. These internal surface dimensions of the shellcan be obtained by either of the above methods immediately after the hotsizing operation or after any machining operation subsequent to the hotsizing operation.

FIG. 19 shows the numerically controlled machining of the matingablative liner section 50 by a tape controlled machining head 46. Thisnumerically controlled machining technique reproduces the size and shapeof the internal surfaces of the shell 10 in FIG. 18 on the outsidesurfaces of the mating ablative liner component 45 in FIG. 19. Byemploying such a machining technique the mating and fit-up problems areminimized and the ablative liner sections 50, 51 and 52 can be fittedand attached without requiring an excessively thick adhesive line.

Referring to FIG. it is seen that the shell ablative liner material 11,14, 15 and 16 is fitted into and bonded to the shell 10 in varioussections 50, 51 and 52. The ablative liner material 11 comprising thethroat sections 51, is composed of wrapped graphite fabric tape, biascut, and impregnated with a phenolic resin which is cured by acombination of pressure and temperature. After the curing cycle theablative liner material 11 is machined and then wrapped with abidirectional glass fabric 12 impregnated with a phenolic resin. Thiscomposite section 51 is then cured by a combination of pressure andtemperature, machined as previously indicated by numerically controlledmachining techniques and bonded to the shell 10. The ablative linermaterial 14 comprising the convergent or inlet cone section 50 is madeup of a carbon fabric tape, bias cut and impregnated with a phenolicresin. This section 50 is also wrapped with the glass fabric 12 and iscured, machined and bonded to the shell 10 in the same manner as issection 151. The divergent or exit cone section 52 is constructed withtwo ablative liner materials 14 and 16. The ablative liner material 16is made up of a high-silica fabric tape, warp cut and impregnated with aphenolic resin. This section 52 is wrapped with the same glass fabric 12as sections 50 and 51 and is also cured, machined and bonded to theshell 10 in the same manner. A reinforcement and stiffening is providedfor the end of the exit cone section 52 by the additional wrapping ofhigh-strength glass rovings 17, which are impregnated with an epoxyresin and cured at room temperature.

The different ablative liner materials 11, 14 and 16 are used because ofthe variation in the design criteria encountered at different locationsthroughout a rocket nozzle. The point of major concern in the inlet conesection 50 design is to insulate the shell 10 from the erosive hightemperature gas flow in this subsonic flow region. The uniformity ofmaterial is very important in this area because of the erraticdownstream erosion which can be caused by turbulent fiow resulting fromexcessive channeling or irregular erosion in the exit cone section 52.The ablative liner material 14 made up of carbon-reinforced bias-cuttape has the proper fiber orientation and density to withstand the hightemperature gas flow while insulating the shell 10. This ablative linermaterial 14 is also capable of predictable erosion characteristics whichare also necessary in the design of the inlet cone section 50.

The throat sections 51 are also composed of bias-cut tape but with agraphite reinforcement. The major concern with the ablative linermaterial 11 in the throat sections 51 is resistance to thermal shock andattendant damage due to cracks and excessive channeling. The use oftape-wrapped graphite throat sections 51 minimizes this problem byproviding a more elastic material, better able to withstand the thermalshock and possessing greater structural integrity due to thecircumferential continuity of the graphite reinforcement fibers.

The ablative liner material 14 used in the forward portion of the exitcone section 52 consists of carbonreinforced tape because of the hightemperatures and severe erosion conditions encountered just aft of thethroat section 51 and because of the necessity for providing gooddownstream support to the throat section 51. In the aft portion of thisexit cone section 52, the ablative material 16 utilized is a lessexpensive silicareinforced tape having a slightly higher erosion ratethan the carbon but adequate for the less severe conditions encounteredin this area.

Attention is now directed to FIG. 21 which shows the completed rocketnozzle 9. After all the ablative liner sections 50, 51 and 52 have beenbonded to the shell 10 and cured, structural tie laminate, bidirectionalglass fabric tape 18, warp cut, and wet-dip coated with a phenolic resinis wrapped around the outside of the shell 10 from the forward end ofthe middle throat section 51 to just beyond the forward portion of theexit cone section 52. This structural tie laminate tape 18 providesgreater support for the shell 10 and the exit cone section 52 and allowsthe exit cone section 52 to withstand the resulting operational thrustforces. Glass roving 19, Wetdip coated with an epoxy resin is thenwrapped around the outside of the shell 10 at the aft two throatsections 51. This glass roving 19 is added to provide reinforcement ofthe shell required due to an increased pressure buildup at this areaduring nozzle operation. Both the structural tie laminate tape 18 andthe glass roving 19 are cured at room temperature.

From the foregoing it will now be apparent that a novel and uniquerocket nozzle construction and process of manufacture thereof which arewell adapted to fulfill the aforestated objects of the invention. Whilevarious alternative embodiments and methods which fall within the scopeof the present invention may suggest themselves of those skilled in theart, it is intended in the appended claims to cover all such additionalembodiments, constructions and methods which fall within the spirit andscope of the invention.

Having thus described the invention, what is claimed as new and usefuland what is desired to be secured by Letters Patent is:

1. A process for manufacturing an integrated rocket nozzle shellcomprising the steps of cutting flat plate into a plurality of platesegments, contour forming said plate segments into arcuatefrusto-conical ring segments, Welding said segments into a plurality ofdiscrete frustoconical ring sections of varying diameters and coneangles, hot sizing said ring sections to remove distortion, welding thering sections together to form a unitary shell structure, hot sizing theunitary shell structure to remove distortion, and machining the externalsurfaces of the shell structure.

2. A process for manufacturing an integrated rocket nozzle shell as inclaim 1, said hot sizing of the ring section and shell parts comprisingthe steps of heating the part to the annealing temperature of thematerial, placing the heated part over a rigid fixture having a surfaceconforming to the desired dimensions of the part, and imparting a forceto the part to drive said part against said rigid sizing fixture therebyconforming to the size and shape of the said part to that of thefixture.

3. A process for manufacturing a rocket nozzle-ablative liner compositecomprising the steps of measuring and recording the dimensions of theinternal surfaces of a nozzle shell made in the process of claim 1,match machining the external surface dimensions of the ablative linersections to correspond with the internal surfaces of the nozzle shell,and interfitting and attaching the match machined ablative linersections to the 110Zzle shell.

4. A process of match machining, interfitting and attaching ablativeliner sections to a nozzle shell made in the process of claim 1comprising the steps of measuring and recording by numerical control theinternal surface surfaces of the mating ablative liner sections bynumerical control using said recorded dimensions, and interfitting andattaching said liner sections to the nozzle shell.

References Cited UNITED STATES PATENTS 1,621,007 3/1927 Ford 29-47l.12,337,049 12/ 1943 Jackson 29475 2,700,988 2/1955 Smisko 29463 2,732,3231/1956 Linnert 29487 X 2,948,061 8/1960 Carstens 29463 3,052,021 9/1962Needham 29493 X 3,082,601 3/1963 Witt.

3,184,362 5/1965 Litsky et al. 29493 X 3,186,063 6/1965 Dopp 29-475 XCHARLIE T. MOON, Primary Examiner.

PAUL M. COHEN, Assistant Examiner.

v us. c1. X.R. 29 407, 463, 471.1, 475, 487, 493, 497

